Active suspension method and apparatus for a rotary wing aircraft

ABSTRACT

In a method of controlling the hardness of a damper in an aircraft landing gear, the aircraft having a structure that is deformable at a natural frequency, the hardness of the damper is adjusted during a contraction of the damper as a function of the natural frequency.

The present invention relates to an active suspension method andapparatus for a rotary wing aircraft.

The technical field of the invention is that of manufacturing suspensionsystems for helicopters.

The present invention applies to aircraft landing gear, in particularfor rotorcraft, serving to improve the performance thereof in the eventof a hard landing or a crash landing.

Various systems have already been described for adapting the stiffnessand/or the hardness of damping in a vehicle suspension.

U.S. Pat. No. 4,730,816 describes an apparatus for controlling thehardness of a suspension for a road vehicle; the apparatus has twochambers interconnected by a pipe fitted with an electromagneticallycontrolled valve having two constrictions, and a hydropneumaticaccumulator acting as a spring; the single valve is fast, and itinfluences the hardness both of the spring and of the damper as afunction of signals delivered by sensors to an electronic apparatus thatcontrols the position of the valve. The hardness of damping is measuredat all times and its desired value is calculated by the electronicapparatus, which value is then applied to the control valve; thehardness of damping and of the suspension can vary as a function of thestate of the road, the state of the load, the acceleration of thevehicle, or the speed of the vehicle, e.g. in application of aservo-control relationship that reduces the hardness of the suspensionat slow speed.

U.S. Pat. No. 5,276,622 describes a system for controlling variation inthe damping coefficient of a damper for the purpose of avoidingend-of-stroke knocks.

U.S. Pat. No. 6,120,009 describes a shock absorber for an aircraftlanding gear of length, stiffness, and damping hardness that arevariable as a function of signals delivered by sensors to an electronicapparatus controlling three valves that control the flow rate of fluidentering or leaving one of the chambers of the shock absorber; thesensors may be sensitive to operating parameters of the damper such asthe speed of its piston, to operating parameters of the airplane such asits speed, and/or to parameters of the landing gear.

The document “Testing of semiactive landing gear control for generalaviation aircraft” by Gian Luca Ghiringhelli, Journal of Aircraft, Vol.37, No. 4, July-August 2000, pp. 606-616, proposes mathematical modelsfor the behavior of landing gear tires and dampers, and describesexperiments for validating those models and their applications tocontrolling a valve for controlling the hardness of a damper.

U.S. Pat. No. 6,474,598 describes an oleopneumatic damper for landinggear including an electromagnetic coil for controlling the viscosity ofa magneto-rheological oil flowing from one chamber of the damper to theother.

An object of the invention is to propose an improved active suspensionsystem for an aircraft; an object of the invention is to remedy thedrawbacks and shortcomings of prior art aircraft suspension systems.

An object of the invention is to provide an improved method andapparatus for the suspension of a rotary wing aircraft, and inparticular for a helicopter.

An object of the invention is to propose helicopter landing gear withservo-controlled damping adapted to the helicopter, and/or adapted tohard landings (and/or to crash landings).

An object of the invention is to propose an aircraft fitted with animproved system for servo-controlling the hardness of landing gear.

The invention applies in particular to helicopters having an airframe orcentral structure supporting a main rotor and at least one engine unit;the helicopter also comprising a tail boom connected to the centralstructure and supporting a tail rotor.

In accordance with an aspect of the invention, the hardness of a damperof landing gear of the aircraft is servo-controlled as a function—inparticular—of data representative of the period corresponding to thenatural frequency of deformation of the structure of the aircraft.

In other words, and in accordance with another aspect of the invention,a method of controlling the hardness of an aircraft landing gear damperis proposed, for an aircraft having a structure that is deformable witha natural mode of deformation and at a natural frequency, in whichmethod said hardness is adjusted during a contraction (or compression)of the damper, as a function of said natural frequency.

According to preferred embodiments:

-   -   a force transmitted by the landing gear to the deformable        structure is measured, and said hardness is adjusted as a        function of the measured transmitted force;    -   alternatively—or in addition—it is possible to measure a rate of        compression of the damper, and to adjust said hardness as a        function of the measured rate of compression;    -   for a damper presenting two fluid chambers communicating via a        duct fitted with a valve of variable aperture, the aperture of        the valve is caused to vary so as to vary the hardness of the        damper;    -   the hardness of the damper is adjusted to spread out the energy        of a landing impact over a duration close to the reciprocal of        said natural frequency;    -   a servo-control relationship for opening a variable aperture        valve of the damper is selected and/or generated as a function        of parameters for the current landing; and    -   in order to select or generate said relationship, use is made of        recorded and/or measured data relating to the trim of the        aircraft, the vertical speed of the aircraft, the lift of the        aircraft, the mass of the aircraft, the position of the center        of gravity of the aircraft, the position of the landing gear of        the aircraft, the stiffness of the landing gear of the aircraft,        and the drive linkage of the landing gear of the aircraft.

In other words, and according to another aspect of the invention,apparatus is proposed for servo-controlling the hardness of a landinggear damper of an aircraft having a structure that presents at least onenatural frequency of deformation; the apparatus includes a system foradjusting said hardness and/or a force transmitted by the landing gearto the structure—during landing—as a function of duration datacorresponding to the reciprocal of said natural frequency.

In a preferred embodiment, while landing the aircraft and during thecompression (shortening) stroke of a damper in a landing gear connectedto the structure of the aircraft, the hardness of the damper isservo-controlled so that the force transmitted by the landing gear tothe structure of the aircraft is at a maximum for a duration close tothe period corresponding to a natural frequency of deformation of thestructure of the aircraft.

Advantageously, the hardness of the damper is servo-controlled so thatthe ratio of the duration (D) to the period (P) is greater than 0.8 andless than 1.2; which is equivalent to-the product of said durationmultiplied by the natural frequency in question lying in the same rangeof values.

Ideally, the hardness of the damper is servo-controlled so that theratio of the duration (D) to the period (P) is equal to 1; which isequivalent to the product of said duration multiplied by the naturalfrequency in question being equal to 1.

In the particular circumstance of a rotary wing aircraft of structurehaving a natural mode of deformation in bending (laterally orvertically) for the tail boom, the hardness of the damper is controlled,during landing at a vertical speed of approach of not less than 2 metersper second, in such a manner as to spread out the energy and/or theforce transmitted by the landing gear to the structure in time over aduration close to the period calculated from the natural frequencycorresponding to said natural mode of deformation of the tail boom.

The invention makes it possible to minimize the dynamic response of thestructure during a landing by actively controlling the behavior of thelanding gear.

The invention thus makes it possible to limit the peak forces exerted onthe structure and to limit the deformations that result therefrom; undercertain conditions, this decrease can reach or even exceed 40%, therebyenabling the structure to be of smaller dimensions and consequentlyreducing its weight, or else enabling the aircraft to have improvedability to cope with a hard landing and/or a crash landing, i.e. toincrease its maximum admissible impact speed.

Other characteristics and advantages of the invention appear from thefollowing description which refers to the accompanying drawing thatshows preferred embodiments of the invention, without any limitingcharacter.

FIG. 1 is a diagram of an active control system for an undercarriage inaccordance with the invention.

FIG. 2 shows an example of how a sensor for measuring the forcetransmitted to the structure by an undercarriage can be located.

FIG. 3 is a diagrammatic timing chart showing the matching of theduration of a force “pulse” transmitted by the undercarriage, to theinverse of the natural frequency of the structure excited by theundercarriage.

The invention is based in particular on studies and tests performed onhelicopters that have revealed coupling between the structure of thehelicopter and its undercarriage, in particular when landing at a highvertical speed (2 meters per second (m/s)<speed<6 m/s). Under certainconditions, it appears in particular that the reaction exerted by theground on the tire of an undercarriage and as transmitted by theundercarriage (including its oleo-strut) to the structure of thehelicopter, can excite one of the natural deformation modes of thestructure.

In the particular case of a helicopter, the tail boom is a structuralelement that is heavily stressed during landing because of its highlycantilevered configuration relative to the central structure; the momentof the inertia forces at its connection is high.

Stress recordings have shown not only that the tail boom is loadedstatically by the sudden rise in force delivered by the landing gear,but that it is also loaded dynamically by a resonance phenomenon, at anatural frequency corresponding to a natural mode of deformation inbending of the tail boom.

The servo-control system of the invention serves to match the excitationof the structure by the landing gear to the natural frequency of themost critical structural element in such a manner as to minimize thepeak value of the force in said structural element, thus making itpossible for it to be more lightly dimensioned.

Excitation of the structure by the landing gear is controlled byspreading out the width of the force pulse transmitted by the landinggear to a greater or lesser extent.

The diagram of FIG. 3 plots variation in the excitation EF of thestructure by the landing gear up the ordinate as a function of timealong the abscissa, and also plots the response RS of the structure tosaid excitation.

Writing the duration of the excitation signal as D and the angularfrequency of the structural response as ω, the purpose of theservo-control relationship is to determine the value Dmin of theduration D that minimizes the structural response RS; the minimumresponse is generally obtained when the ratio τ=D/(2π/ω) is equal to 1,which corresponds to the signals being in phase opposition: theexcitation signal EF becomes zero simultaneously with the structuralresponse, thereby reducing said response. Conversely, a maximum responseis obtained specifically when the ratio is equal to 1/2 or 3/2: undersuch circumstances, the falling edge of the excitation relaunches thestructural response.

By way of example, for a deformable helicopter structure presenting anatural frequency of 5 hertz (Hz), the opening of the valve isservo-controlled so as to obtain a force pulse exerted on the structurehaving a duration close to 200 milliseconds (ms) e.g. situated in therange 160 ms to 240 ms.

The simplified expression for the landing gear force as a function oftime is of the form:EF(t)=C(t).dx/dtwhere C(t) designates the damping coefficient of the landing gear,having a value that is adjusted at each instant by the servo-controlsystem, and where dx/dt designates the rate at which the landing gear islengthening or shortening.

Given that the force EF is bounded by the static strength limit EFmax ofthe connection between the landing gear fitting and the structure of theaircraft, i.e. the point where force is injected into the structure, andgiven that speed is measured continuously by means of a sensor, theservo-control computer seeks in real time the value C(t) that will givethe ratio T that is the closest to 1 while still complying with theconstraint EF<EFmax.

In a variant, the expression for EF(t) can include additional terms fortaking account specifically of the stiffness and damping characteristicsof the tire and of the gas in the landing gear damper. The generalexpression for EF(t) then has the following form;EF(t)=C(t).dx/dt+Kxwhere K is a stiffness term (tire, gas, . . .)

The system of the invention for active control of landing gear serves toservo-control the curve of the force EF transmitted by the landing gearto the conditions of landing by using a preestablished relationshipadapted to the particular conditions of the present landing; therelationship is devised to limit the peak value of the response RS ofthe deformable structure, and to spread out the energy of the impactover time.

To implement such active control, it is preferred to have the meansshown in FIG. 1, i.e.:

a) a memory 34 for storing memorized and/or measured data defining the“landing configuration”: landing conditions (trim, vertical speed,hovering, . . .), mass and center of gravity of the aircraft,characteristics of the landing gear (position, stiffness, linkage, . ..);

b) a sensor 22 for measuring the forces exerted by the landing gear onthe structure of the helicopter during landing, and/or a sensorresponsive to the compression rate of the damper;

c) a computer 31, 32, and 33 serving to select or to generate arelationship (force =f(t)) satisfying the condition EF<EFmax and adaptedto the landing parameters specific to the current flight, and also toservo-control the hydraulic throttling in compliance with saidrelationship; and

d) an actuator, driven by the computer, for the purpose of modulatingthe damping characteristics of the landing gear.

When the servo-control relationships are predetermined, they are storedin the memory 34; under such circumstances, a selector 33 in thecomputer 31 responds to information delivered thereto by the memory 34,and selects the relationship that is best adapted to the landingparameters of the moment.

In a variant, a servo-control relationship may be generated by a module(not shown) for computing the servo-control relationship and that isimplanted in the computer.

A servo-control module 32 may serve to control the actuator 18 on thebasis of force measurements delivered by a sensor 22, while complyingwith the selected servo-control relationship.

The actuator includes a control member acting directly on the dampingcharacteristics of the landing gear (by throttling the fluid), and meansfor driving said control member under the control of the computerserving to amplify the electrical control signal and to convert it intomechanical movement of the above-mentioned member.

The control member may be an orifice of variable section, of thediaphragm type. However, the stresses internal to the landing gear maymake it difficult to dimension such a member; in which case it may bepreferable to use a member of the valve type, which closes to a greateror less extent an orifice for throttling the flow of oil from onechamber of the damper to another. The member is driven by an electricmotor under the control of the computer.

In a variant, the member may be driven hydraulically since the movementof compressing the landing gear during landing causes the dampinghydraulic fluid to move: this “free” energy source is available; thedrive means is then of the servo-control type, with a distributor valvecontrolled electrically by the computer.

In another variant, instead of modifying the fluid throttling orifice,it is possible to use a fluid whose viscosity characteristics aremodified, in particular a magneto-rheological fluid of viscosity thatcan be controlled by means of an electric current.

The measurement, acquisition, and/or recording—in a memory of thecomputer system 31 to 34 on board the aircraft in question—of theparameters relating to the current landing and of the parametersspecific to the aircraft in question can be implemented in particular byusing the systems described in U.S. Pat. Nos. 4,312,042 and 3,426,586;in order to measure the forces transmitted to the landing gear by theaxle and transmitted to the structure of the aircraft by the landinggear, it is possible to use inclinometers or strain gauges JDC1, JDC2secured upstream and downstream from the damper 10.

The position and/or speed of compression (and conversely of expansion)of an oleopneumatic damper 10 of the landing gear 19 along thelongitudinal axis 16 can be determined from the signals delivered by twoaccelerometers (ACC1 and ACC2, FIG. 2) secured respectively to parts orassemblies interconnected by the damper 10: a part 13 connecting an axle11 (carrying wheels 12) to the bottom end of the damper 10, and alsoparts 14, 15 (shown in part and diagrammatically) connecting the top endof the damper 10 to the main structure of the aircraft (not shown).

1. A method of controlling the hardness of a damper (10) of an aircraftlanding gear (19), the aircraft comprising a structure (14, 15, 21)deformable with a natural mode of deformation at a natural frequency(FP) corresponding to said natural mode of deformation, the method beingcharacterized in that the hardness is adjusted during compression of thedamper as a function of said natural frequency.
 2. A method according toclaim 1, in which a force (EF) transmitted by the landing gear to thedeformable structure is measured, and said hardness is adjusted as afunction of the measured transmitted force.
 3. A method according toclaim 1, in which a rate of compression of the damper is measured, andsaid hardness is adjusted as a function of the measured rate ofcompression.
 4. A method according to claim 1, in which the damperpresents two fluid chambers communicating via a duct fitted with a valve(18) of variable aperture, and in which the aperture of the valve iscaused to vary to adjust the hardness of the damper.
 5. A methodaccording to claim 1, in which the hardness of the damper is adjusted tospread out the energy of a landing impact over a duration (D) close tothe reciprocal (Dmin) of said natural frequency.
 6. A method accordingto claim 1, in which a servo-control relationship for opening a valve ofvariable aperture of the damper is selected and/or generated as afunction of parameters for the current landing.
 7. A method according toclaim 6, in which in order to select or generate said relationship, useis made of recorded and/or measured data related to the trim of theaircraft, the vertical speed of the aircraft, the lift of the aircraft,the mass of the aircraft, the position of the center of gravity of theaircraft, the position of the landing gear of the aircraft, thestiffness of the landing gear of the aircraft, and the drive linkage ofthe landing gear of the aircraft.
 8. Apparatus for servo-controlling thehardness of a damper (10) of landing gear (19) of an aircraft having astructure (14, 15, 21) presenting a natural mode of deformation and anatural frequency of deformation (FP) corresponding to said natural modeof deformation, the apparatus being characterized in that it includes asystem (31 to 34) for adjusting the hardness of the damper duringcompression of the damper as a function of duration data (Dmin)corresponding to the reciprocal of said natural frequency.
 9. Apparatusaccording to claim 8, including a sensor (JDC1) responsive to the forcesexerted on the landing gear, suitable for delivering signals or data tothe adjustment system, which are representative of the forces exerted onthe landing gear.
 10. Apparatus according to claim 8, including a sensor(JDC2, 22) responsive to the forces exerted by the landing gear on thestructure, and suitable for delivering signals or data to the adjustmentsystem, which are representative of the forces exerted by the landinggear.
 11. Apparatus according to claim 8, including a sensor (ACC1,ACC2) responsive to the position or the rate of compression of thelanding gear and suitable for delivering signals or data to theadjustment system, which are representative of the position or the rateof compression of the landing gear.
 12. Apparatus according to claim 8,in which the damper comprises first and second chambers and a variableaperture throttling member (18) for controlling the rate of flow offluid from the first chamber to the second chamber, and in which thesystem for adjusting the hardness of the damper comprises: a memory (34)for storing landing configuration data selected from landing conditiondata, mass data and center of gravity data for the aircraft, and datacharacteristic of the landing gear; at least one sensor (ACC1, ACC2)responsive to the rate of contraction of the damper, the sensordelivering a rate of contraction signal; and data processor means (31,32, 33) coupled to the data storage memory, to the measurement sensor,and to the throttling member, the data processor means serving to selector generate a servo-control relationship for the aperture of thethrottling member adapted to the landing configuration data specific tothe current flight, and also to servo-control the aperture of thethrottling member in application of said relationship and as a functionof the contraction rate signal delivered by the sensor, so as to adjustthe hardness of the damper while it is being compressed, in such amanner that the force transmitted by the landing gear to the structureof the aircraft is at a maximum for a duration (D) close to saidduration data (Dmin).
 13. Apparatus according to claim 12, furthercomprising a sensor (JDC1, JDC2, 22) for measuring the forces exerted bythe landing gear on the structure of the aircraft while landing, thesensor delivering a force signal, and in which said data processor meansis connected to the force measuring sensor and servo-controls thethrottling member also as a function of the force signal.
 14. Apparatusaccording to claim 12, in which said at least one sensor (ACC1, ACC2)responsive to the rate of contraction of the damper comprises twoaccelerometers secured respectively to a part (13) connecting an axle(11) to the damper (10), and to a part (14, 15) of the structure of theaircraft.
 15. Apparatus according to claim 8, in which the damper has afirst chamber and a second chamber and a variable aperture throttlingmember (18) for controlling the flow of a fluid from the first chamberto the second chamber, and in which the system for adjusting thehardness of the damper comprises: a memory (34) for storing landingconfiguration data selected from landing condition data, data concerningthe mass and the center of gravity of the aircraft, and datacharacteristic of the landing gear; a sensor (JDC1, JDC2, 22) formeasuring the forces exerted by the landing gear on the structure of theaircraft during a landing, the sensor delivering a force signal; anddata processor means (31, 32, 33) coupled to the data storage memory, tothe measurement sensor, and to the throttling member, the data processormeans serving to select or generate a servo-control relationship for theaperture of the throttling member adapted to the landing configurationdata specific to the current flight, and also serving to servo-controlthe aperture of the throttling member in application of said generationand as a function of the force signal delivered by the sensor, so as toadjust the hardness of the damper while it is being compressed, toensure that the force transmitted by the landing gear to the structureof the aircraft remains at a maximum for a duration (D) close to saidduration data (Dmin).